Flow ranges for centrifugal compressors generally, and the ranges over which they may operate specifically, are dictated by the stalling characteristics of the compressor impeller and a diffuser which receives compressed air from the impeller. The stalling characteristics in turn are intrinsically controlled by contours of the impeller blades as well as Mach numbers achieved during operation.
Although various centrifugal compressors may utilize either vaned or vaneless diffuser systems, where maximum efficiency at high Mach numbers is required, the use of vaned diffuser systems becomes almost mandatory. This in turn means that the impeller and diffuser must be matched at peak efficiency flow conditions.
In such cases, the vaned diffuser tends to be the flow controlling component in that the overall Mach number occurring therein generally is higher than that of the compressor inducer which operates with a larger variation of Mach numbers over the radius of the blades extending from a hub of the impeller to the compressor shroud in an axial inflow, radial outflow centrifugal compressor. The diffused flow from the impeller in such a case, and non-uniform entrance conditions which result, further aggravate stalling sensitivity. To attain a large flow range requires that the impeller and the diffuser must be capable of operating into "positive incidence" or stalled regions to flows where compressor surge is eventually triggered. Compressor surge is generally believed to stem from operation on an unstable portion of the overall compressor characteristic (a positive slope portion) where the impeller static pressure ratio decreases with decreasing flow. Thus, one effective method of increasing compressor operating range is to provide sufficient impeller stability so that the downstream diffuser can operate well into its positive incidence zone, even though the diffuser static pressure recovery versus flow characteristic exhibits a positive slope. Impeller stability is conventionally provided by the use of blade tip backsweep since the increased backsweep provides a more negative sloped static pressure rise versus flow characteristic. However, increasing the tip backsweep increases stresses appearing in the impeller blade and/or hub.
Present-day advanced aircraft require auxiliary power units (APUs) as a supply of electrical, hydraulic, and pneumatic power to secondary power systems of the aircraft. Generally speaking, the APUs are gas turbine units and must be highly reliable. In addition, compactness is also required. Most suitably, the APUs are then based upon a single shaft, constant speed, gas turbine having a high specific speed, single stage centrifugal compressor, a reverse flow annular combustor, and a single stage radial or axial inflow turbine. Shaft power is utilized to drive electrical generators and/or pumps and compressor bleed air extracted from the system prior to combustion to provide pneumatic power. For high bleed air output, it is necessary to design the compressor to operate adjacent to its maximum flow point, that is, near a so-called "choke" condition. The extraction of increasing amounts of shaft power at constant speed and constant turbine inlet temperatures from the choke point incrementally displace the compressor operating point to lower flows and are eventually limited by encroachment upon the compressor surge line.
The assignee of the instant invention has in the past embraced the problems as defined herein before and defined a solution as described in now issued U.S. Pat. No. 4,981,018. This solution called for a compressor construction wherein a compressor hub and associated blades of the impeller formed thereby are surrounded by an annular shroud. Bleed passages in the shroud which are angled in the direction of flow, both axially and radially, have been found to improve efficiency.
While the invention of U.S. Pat. No. 4,981,018 advanced the state of the art, low operating temperature environmental operating conditions conspired to detract from the turbomachine's overall efficiency.
The shroud bleed air which has been warmed during compression is reingested by the compressor at the compressor inlet. This warmed bleed air appears to stratify toward a compressor outer shroud wall. This warmed stratified bleed air enters an outer flow path of the compressor's inducer section and causes performance penalties to the turbomachine.
In cold ambient air during startup, cold air is drawn into the compressor inlet and in so doing cools inlet support struts and related structure in such a manner as to cause the compressor front shroud structure to pull away from the compressor rotor which further deteriorates the turbomachine's performance.
The invention to be described more fully hereinafter solves the problems just enumerated in an exceptionally simple manner.